Ice reduction mechanism for turbofan engine

ABSTRACT

A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.

TECHNICAL FIELD

The present subject matter relates generally to a gas turbine engine, ormore particularly to a gas turbine engine configured to reduce icebuildup or ice formation on inlet components of the engine.

BACKGROUND

A turbofan engine generally includes a fan having a plurality of fanblades and a turbomachine arranged in flow communication with oneanother. Additionally, the turbomachine of the turbofan engine generallyincludes, in serial flow order, a compressor section, a combustionsection, a turbine section, and an exhaust section. In operation, air isprovided from the fan to an inlet of the compressor section where one ormore axial compressors progressively compress the air until it reachesthe combustion section. Fuel is mixed with the compressed air and burnedwithin the combustion section to provide combustion gases. Thecombustion gases are routed from the combustion section to the turbinesection. The flow of combustion gasses through the turbine sectiondrives the turbine section and is then routed through the exhaustsection, e.g., to atmosphere.

However, during inclement weather, freezing rain, hail, sleet, ice,etc., can accumulate on the inlet components of the turbofan engine.When ice accumulates, it can break off and be ingested into the engine.Further, large portions of ice can damage the fan blades or otherdownstream components of the engine, and may potentially cause an engineflameout.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to an exemplary embodiment of the present subjectmatter.

FIG. 2 is a close-up, schematic, cross-sectional view of a forward endof the exemplary gas turbine engine of FIG. 1 according to an exemplaryembodiment of the present subject matter.

FIG. 3 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to an exemplary embodiment of the present subjectmatter.

FIG. 4 it is a schematic view of an inlet to a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

FIG. 5A is a schematic cross-sectional view of an exemplary gas turbineengine according to an exemplary embodiment of the present subjectmatter.

FIG. 5B is a cross-sectional view of a part span inlet guide vane of anexemplary gas turbine engine at a first location along a span of thepart span inlet guide vane according to an exemplary embodiment of thepresent subject matter.

FIG. 6 is a schematic cross-sectional view of an exemplary gas turbineengine according to another exemplary embodiment of the present subjectmatter.

FIG. 7 is a schematic cross-sectional view of an exemplary gas turbineengine according to another exemplary embodiment of the present subjectmatter.

FIG. 8A is a schematic cross-sectional view of an exemplary gas turbineengine according to another exemplary embodiment of the present subjectmatter.

FIG. 8B is a cross-sectional view of a part span inlet guide vane of anexemplary gas turbine engine at a first location along a span of thepart span inlet guide vane according to another exemplary embodiment ofthe present subject matter.

FIG. 9A is a schematic cross-sectional view of an exemplary gas turbineengine according to another exemplary embodiment of the present subjectmatter.

FIG. 9B is a cross-sectional view of a part span inlet guide vane of theexemplary gas turbine engine of FIG. 9A at a first location along a spanof the part span inlet guide vane according to another exemplaryembodiment of the present subject matter.

Corresponding reference characters indicate corresponding partsthroughout the several views. The exemplifications set out hereinillustrate exemplary embodiments of the disclosure, and suchexemplifications are not to be construed as limiting the scope of thedisclosure in any manner.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

The following description is provided to enable those skilled in the artto make and use the described embodiments contemplated for carrying outthe disclosure. Various modifications, equivalents, variations, andalternatives, however, will remain readily apparent to those skilled inthe art. Any and all such modifications, variations, equivalents, andalternatives are intended to fall within the scope of the presentdisclosure.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

For purposes of the description hereinafter, the terms “upper”, “lower”,“right”, “left”, “vertical”, “horizontal”, “top”, “bottom”, “lateral”,“longitudinal”, and derivatives thereof shall relate to the disclosureas it is oriented in the drawing figures. However, it is to beunderstood that the disclosure may assume various alternativevariations, except where expressly specified to the contrary. It is alsoto be understood that the specific devices illustrated in the attacheddrawings, and described in the following specification, are simplyexemplary embodiments of the disclosure. Hence, specific dimensions andother physical characteristics related to the embodiments disclosedherein are not to be considered as limiting.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine, with forward referring to a position closer to an engineinlet and aft referring to a position closer to an engine nozzle orexhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Additionally, the terms “low,” “high,” or their respective comparativedegrees (e.g., lower, higher, where applicable) each refer to relativespeeds or pressures within an engine, unless otherwise specified. Forexample, a “low-pressure turbine” operates at a pressure generally lowerthan a “high-pressure turbine.” Alternatively, unless otherwisespecified, the aforementioned terms may be understood in theirsuperlative degree. For example, a “low-pressure turbine” may refer tothe lowest maximum pressure turbine within a turbine section, and a“high-pressure turbine” may refer to the highest maximum pressureturbine within the turbine section. An engine of the present disclosuremay also include an intermediate pressure turbine, e.g., an enginehaving three spools.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin. These approximating margins may apply to asingle value, either or both endpoints defining numerical ranges, and/orthe margin for ranges between endpoints.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

As used herein, the term “fan pressure ratio” refers to a ratio of anair pressure immediately downstream of the fan blades if a fan duringoperation of the fan to an air pressure immediately upstream of the fanblades of the fan during operation of the fan.

As used herein, the term “rated speed” with reference to a turbofanengine refers to a maximum rotational speed that the turbofan engine mayachieve while operating properly. For example, the turbofan engine maybe operating at the rated speed during maximum load operations, such asduring takeoff operations.

Also as used herein, the term “fan tip speed” as defined by theplurality of fan blades of the fan refers to a linear speed of an outertip of a fan blade along a radial direction during operation of the fan.

The present disclosure is generally related to a means for reducing icebuildup or ice formation on inlet components of an engine, e.g., aninlet pre-swirl feature configured as a plurality of part span inletguide vanes.

In some exemplary embodiments of the present disclosure, the means forreducing ice buildup or ice formation includes a heat source that is inthermal communication with the inlet pre-swirl feature, e.g., configuredas a plurality of part span inlet guide vanes. For example, in a firstexemplary embodiment, the heat source includes an engine bleed airflow.In another exemplary embodiment, the heat source includes an electricalheating element disposed in a leading edge of the inlet pre-swirlfeature. In yet another exemplary embodiment, the heat source includesan engine oil.

In other exemplary embodiment, the means for reducing ice buildup or iceformation includes a vibration assembly in mechanical communication withthe inlet pre-swirl feature, e.g., configured as a plurality of partspan inlet guide vanes, or an anti-icing coating on the pre-swirlfeature.

Inclusion of one or more of these means for reducing ice buildup or iceformation provides an anti-icing or de-icing mechanism that prevents thebuildup and shedding of pieces of ice into the engine during, e.g.,adverse weather conditions, resulting in safer operation of the gasturbine engine.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is an aeronautical,turbofan jet engine 10, referred to herein as “turbofan engine 10”,configured to be mounted to an aircraft, such as in an under-wingconfiguration or tail-mounted configuration. As shown in FIG. 1 , theturbofan engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference), a radial directionR, and a circumferential direction (i.e., a direction extending aboutthe axial direction A; not depicted). In general, the turbofan 10includes a fan section 14 and a turbomachine 16 disposed downstream fromthe fan section 14 (the turbomachine 16 sometimes also, oralternatively, referred to as a “core turbine engine”).

The exemplary turbomachine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a first, booster or low pressure (LP) compressor 22and a second, high pressure (HP) compressor 24; a combustion section 26;a turbine section including a first, high pressure (HP) turbine 28 and asecond, low pressure (LP) turbine 30; and a jet exhaust nozzle section32. A high pressure (HP) shaft 34 drivingly connects the HP turbine 28to the HP compressor 24. A low pressure (LP) shaft 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. The compressor section,combustion section 26, turbine section, and jet exhaust nozzle section32 are arranged in serial flow order and together define a core airflowpath 37 through the turbomachine 16. It is also contemplated thatthe present disclosure is compatible with an engine having anintermediate pressure turbine, e.g., an engine having three spools.

Referring still the embodiment of FIG. 1 , the fan section 14 includes avariable pitch, single stage fan 38, the turbomachine 16 operablycoupled to the fan 38 for driving the fan 38. The fan 38 includes aplurality of rotatable fan blades 40 coupled to a disk 42 in a spacedapart manner. As depicted, the fan blades 40 extend outwardly from disk42 generally along the radial direction R. Each fan blade 40 isrotatable relative to the disk 42 about a pitch axis P by virtue of thefan blades 40 being operatively coupled to a suitable actuation member44 configured to collectively vary the pitch of the fan blades 40, e.g.,in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal centerline 12 by LP shaft 36across a power gear box 46. The power gear box 46 includes a pluralityof gears for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed. Accordingly, for the embodimentdepicted, the turbomachine 16 is operably coupled to the fan 38 throughthe power gear box 46.

In exemplary embodiments, the fan section 14 includes twenty-two (22) orfewer fan blades 40. In certain exemplary embodiments, the fan section14 includes twenty (20) or fewer fan blades 40. In certain exemplaryembodiments, the fan section 14 includes eighteen (18) or fewer fanblades 40. In certain exemplary embodiments, the fan section 14 includessixteen (16) or fewer fan blades 40. In certain exemplary embodiments,it is contemplated that the fan section 14 includes other number of fanblades 40 for a particular application.

During operation of the turbofan engine 10, the fan 38 defines a fanpressure ratio and the plurality of fan blades 40 each define a fan tipspeed. The exemplary turbofan engine 10 depicted defines a relativelyhigh fan tip speed and relatively low fan pressure ratio duringoperation of the turbofan engine at a rated speed. As used herein, theterm “fan pressure ratio” refers to a ratio of an air pressureimmediately downstream of the fan blades 40 during operation of the fan38 to an air pressure immediately upstream of the fan blades 40 duringoperation of the fan 38. For the embodiment depicted in FIG. 1 , the fan38 of the turbofan engine 10 defines a relatively low fan pressureratio. For example, the turbofan engine 10 depicted defines a fanpressure ratio less than or equal to about 1.5. For example, in certainexemplary embodiments, the turbofan engine 10 may define a fan pressureratio less than or equal to about 1.4. The fan pressure ratio may be thefan pressure ratio of the fan 38 during operation of the turbofan engine10, such as during operation of the turbofan engine 10 at a rated speed.

As used herein, the term “rated speed” with reference to the turbofanengine 10 refers to a maximum rotational speed that the turbofan engine10 may achieve while operating properly. For example, the turbofanengine 10 may be operating at the rated speed during maximum loadoperations, such as during takeoff operations.

Also as used herein, the term “fan tip speed” defined by the pluralityof fan blades 40 refers to a linear speed of an outer tip of a fan blade40 along the radial direction R during operation of the fan 38. Inexemplary embodiments, the turbofan engine 10 of the present disclosurecauses the fan blades 40 of the fan 38 to rotate at a relatively highrotational speed. For example, during operation of the turbofan engineat the rated speed, the fan tip speed of each of the plurality of fanblades 40 is greater than or equal to 1000 feet per second and less thanor equal to 2250 feet per second. In certain exemplary embodiments,during operation of the turbofan engine at the rated speed, the fan tipspeed of each of the fan blades 40 may be greater than or equal to 1,250feet per second and less than or equal to 2250 feet per second. Incertain exemplary embodiments, during operation of the turbofan engine10 at the rated speed, the fan tip speed of each of the fan blades 40may be greater than or equal to about 1,350 feet per second, such asgreater than about 1,450 feet per second, such as greater than about1,550 feet per second, and less than or equal to 2250 feet per second.

Referring still to the exemplary embodiment of FIG. 1 , the disk 42 iscovered by rotatable front nacelle or hub 48 aerodynamically contouredto promote an airflow through the plurality of fan blades 40.Additionally, the exemplary fan section 14 includes an annular fancasing or outer nacelle 50 that at least partially, and for theembodiment depicted, circumferentially, surrounds the fan 38 and atleast a portion of the turbomachine 16.

More specifically, the outer nacelle 50 includes an inner wall 52 and adownstream section 54 of the inner wall 52 of the outer nacelle 50extends over an outer portion of the turbomachine 16 so as to define abypass airflow passage 56 therebetween. Additionally, for the embodimentdepicted, the outer nacelle 50 is supported relative to the turbomachine16 by a plurality of circumferentially spaced outlet guide vanes 55.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan engine 10 through an associated inlet 60 of the outernacelle 50 and/or fan section 14. As the volume of air 58 passes acrossthe fan blades 40, a first portion of the air 58 as indicated by arrows62 is directed or routed into the bypass airflow passage 56 and a secondportion of the air 58 as indicated by arrow 64 is directed or routedinto the core air flowpath 37. The ratio between an amount of airflowthrough the bypass airflow passage 56 (i.e., the first portion of airindicated by arrows 62) to an amount of airflow through the core airflowpath 37 (i.e., the second portion of air indicated by arrows 64) isknown as a bypass ratio.

In exemplary embodiments, the bypass ratio during operation of theturbofan engine 10 (e.g., at a rated speed) is less than or equal toabout eleven (11). For example, the bypass ratio during operation of theturbofan engine 10 (e.g., at a rated speed) may be less than or equal toabout ten (10), such as less than or equal to about nine (9).Additionally, the bypass ratio may be at least about two (2).

In other exemplary embodiments, the bypass ratio may generally bebetween about 7:1 and about 20:1, such as between about 10:1 and about18:1. The pressure of the second portion of air indicated by arrows 64is then increased as it is routed through the high pressure (HP)compressor 24 and into the combustion section 26, where it is mixed withfuel and burned to provide combustion gases 66.

In exemplary embodiments, a gear ratio of the power gear box 46 isgreater than or equal to 1.2 and less than or equal to 3.0. In someexemplary embodiments, the gear ratio of the power gear box 46 isgreater than or equal to 1.2 and less than or equal to 2.6. In otherexemplary embodiments, the gear ratio of the power gear box 46 isgreater than or equal to 1.2 and less than or equal to 2.0.

Furthermore, the turbofan engine of the present disclosure also providespre-swirling flow forward of the fan blade tip as described herein.

Referring still to FIG. 1 , the compressed second portion of airindicated by arrows 64 from the compressor section mixes with fuel andis burned within the combustion section to provide combustion gases 66.The combustion gases 66 are routed from the combustion section 26,through the HP turbine 28 where a portion of thermal and/or kineticenergy from the combustion gases 66 is extracted via sequential stagesof HP turbine stator vanes 68 that are coupled to the outer casing 18and HP turbine rotor blades 70 that are coupled to the HP shaft 34, thuscausing the HP shaft 34 to rotate, thereby supporting operation of theHP compressor 24. The combustion gases 66 are then routed through the LPturbine 30 where a second portion of thermal and kinetic energy isextracted from the combustion gases 66 via sequential stages of LPturbine stator vanes 72 that are coupled to the outer casing 18 and LPturbine rotor blades 74 that are coupled to the LP shaft 36, thuscausing the LP shaft 36 to rotate, thereby supporting operation of theLP compressor 22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the turbomachine 16 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air indicated byarrows 62 is substantially increased as the first portion of air 62 isrouted through the bypass airflow passage 56 before it is exhausted froma fan nozzle exhaust section 76 of the turbofan 10, also providingpropulsive thrust. The HP turbine 28, the LP turbine 30, and the jetexhaust nozzle section 32 at least partially define a hot gas path 78for routing the combustion gases 66 through the turbomachine 16.

Referring still to FIG. 1 , the turbofan engine 10 additionally includesa means for reducing ice buildup or ice formation 200 on an inletpre-swirl feature, e.g., configured as a plurality of part span inletguide vanes 100, as described in greater detail herein.

In some exemplary embodiments, it will be appreciated that the exemplaryturbofan engine 10 of the present disclosure may be a relatively largepower class turbofan engine 10. Accordingly, when operated at the ratedspeed, the turbofan engine 10 may be configured to generate a relativelylarge amount of thrust. More specifically, when operated at the ratedspeed, the turbofan engine 10 may be configured to generate at leastabout 20,000 pounds of thrust, such as at least about pounds of thrust,such as at least about 30,000 pounds of thrust, and up to, e.g., about150,000 pounds of thrust. Accordingly, the turbofan engine 10 may bereferred to as a relatively large power class gas turbine engine.

Moreover, it should be appreciated that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 10 may have any othersuitable configuration. For example, in certain exemplary embodiments,the fan may not be a variable pitch fan, the engine may not include areduction gearbox (e.g., power gearbox 46) driving the fan, may includeany other suitable number or arrangement of shafts, spools, compressors,turbines, etc.

As discussed above, the turbofan engine 10 of the present disclosurealso provides pre-swirling flow forward a tip of the fan blade 40.Referring now also to FIG. 2 , a close-up, cross-sectional view of thefan section 14 and forward end of the turbomachine 16 of the exemplaryturbofan engine 10 of FIG. 1 is provided. In exemplary embodiments, theturbofan engine 10 includes an inlet pre-swirl feature located upstreamof the plurality of fan blades 40 of the fan 38 and attached to orintegrated into the outer nacelle 50. More specifically, for theembodiment of FIGS. 1 and 2 , the inlet pre-swirl feature is configuredas a plurality of part span inlet guide vanes 100. The plurality of partspan inlet guide vanes 100 are each cantilevered from of the outernacelle 50 (such as from the inner wall 52 of the outer nacelle 50) at alocation forward of the plurality of fan blades 40 of the fan 38 alongthe axial direction A and aft of the inlet 60 of the outer nacelle 50.More specifically, each of the plurality of part span inlet guide vanes100 define an outer end 102 along the radial direction R, and areattached to/connected to the outer nacelle 50 at the radially outer end102 through a suitable connection means (not shown). For example, eachof the plurality of part span inlet guide vanes 100 may be bolted to theinner wall 52 of the outer nacelle 50 at the outer end 102, welded tothe inner wall 52 of the outer nacelle 50 at the outer end 102, orattached to the outer nacelle 50 in any other suitable manner at theouter end 102.

Further, for the embodiment depicted, the plurality of part span inletguide vanes 100 extend generally along the radial direction R from theouter end 102 to an inner end 104 (i.e., an inner end 104 along theradial direction R). Moreover, as will be appreciated, for theembodiment depicted, each of the plurality of part span inlet guidevanes 100 are unconnected with an adjacent part span inlet guide vane100 at the respective inner ends 104 (i.e., adjacent part span inletguide vanes 100 do not contact one another at the radially inner ends104, and do not include any intermediate connection members at theradially inner ends 104, such as a connection ring, strut, etc.). Morespecifically, for the embodiment depicted, each part span inlet guidevane 100 is completely supported by a connection to the outer nacelle 50at the respective outer end 102 (and not through any structureextending, e.g., between adjacent part span inlet guide vanes 100 at alocation inward of the outer end 102 along the radial direction R). Aswill be discussed below, such may reduce an amount of turbulencegenerated by the part span inlet guide vanes 100.

Moreover, is depicted, each of the plurality of part span inlet guidevanes 100 do not extend completely between the outer nacelle 50 and,e.g., the hub 48 of the turbofan engine 10. More specifically, for theembodiment depicted, each of the plurality of inlet guide vane define aninlet guide vane (“IGV”) span 106 along the radial direction R, andfurther each of the plurality of part span inlet guide vanes 100 furtherdefine a leading edge 108 and a trailing edge 110. The IGV span 106refers to a measure along the radial direction R between the outer end102 and the inner end 104 of the part span inlet guide vane 100 at theleading edge 108 of the part span inlet guide vane 100. Similarly, itwill be appreciated, that the plurality of fan blades 40 of the fan 38define a fan blade span 112 along the radial direction R. Morespecifically, each of the plurality of fan blades 40 of the fan 38 alsodefines a leading edge 114 and a trailing edge 116, and the fan bladespan 112 refers to a measure along the radial direction R between aradially outer tip and a base of the fan blade 40 at the leading edge114 of the respective fan blade 40.

For the embodiment depicted, the IGV span 106 is at least about fivepercent of the fan blade span 112 and up to about fifty-five percent ofthe fan blade span 112. For example, in certain exemplary embodiments,the IGV span 106 may be between about fifteen percent of the fan bladespan 112 and about forty-five percent of the fan blade span 112, such asbetween about thirty percent of the fan blade span 112 and about fortypercent of the fan blade span 112.

Reference will now also be made to FIG. 3 , providing an axial view ofthe inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 . As will beappreciated, for the embodiment depicted, the plurality of part spaninlet guide vanes 100 of the turbofan engine 10 includes a relativelylarge number of part span inlet guide vanes 100. More specifically, forthe embodiment depicted, the plurality of part span inlet guide vanes100 includes between about ten part span inlet guide vanes 100 and aboutfifty part span inlet guide vanes 100. More specifically, for theembodiment depicted, the plurality of part span inlet guide vanes 100includes between about twenty part span inlet guide vanes 100 and aboutforty-five part span inlet guide vanes 100, and more specifically,still, the embodiment depicted includes thirty-two part span inlet guidevanes 100. Additionally, for the embodiment depicted, each of theplurality of part span inlet guide vanes 100 are spaced substantiallyevenly along the circumferential direction C. More specifically, each ofthe plurality of part span inlet guide vanes 100 defines acircumferential spacing 118 with an adjacent part span inlet guide vane100, with the circumferential spacing 118 being substantially equalbetween each adjacent part span inlet guide vane 100.

Although not depicted, in certain exemplary embodiments, the number ofpart span inlet guide vanes 100 may be substantially equal to the numberof fan blades 40 of the fan 38 of the turbofan engine 10. In otherembodiments, however, the number of part span inlet guide vanes 100 maybe greater than the number of fan blades 40 of the fan 38 of theturbofan engine 10, or alternatively, may be less than the number of fanblades 40 of the fan 38 of the turbofan engine 10.

Further, it should be appreciated, that in other exemplary embodiments,the turbofan engine 10 may include any other suitable number of partspan inlet guide vanes 100 and/or circumferential spacing 118 of thepart span inlet guide vanes 100. For example, referring now briefly toFIG. 4 , an axial view of an inlet 60 to a turbofan engine 10 inaccordance with another exemplary embodiment of the present disclosureis provided. For the embodiment of FIG. 4 , the turbofan engine 10includes less than twenty part span inlet guide vanes 100. Morespecifically, for the embodiment of FIG. 4 , the turbofan engine 10includes at least eight part span inlet guide vanes 100, or morespecifically includes exactly eight part span inlet guide vanes 100.Additionally, for the embodiment of FIG. 4 , the plurality of part spaninlet guide vanes 100 are not substantially evenly spaced along thecircumferential direction C. For example, at least certain of theplurality of part span inlet guide vanes 100 define a firstcircumferential spacing 118A, while other of the plurality of part spaninlet guide vanes 100 define a second circumferential spacing 118B. Forthe embodiment depicted, the first circumferential spacing 118A is atleast about twenty percent greater than the second circumferentialspacing 118B, such as at least about twenty-five percent greater such asat least about thirty percent greater, such as up to about two hundredpercent greater. Notably, as will be described in greater detail below,the circumferential spacing 118 refers to a mean circumferential spacingbetween adjacent part span inlet guide vanes 100. The non-uniformcircumferential spacing may, e.g., offset structure upstream of the partspan inlet guide vanes 100.

Referring back to FIG. 2 , it will be appreciated that each of theplurality of part span inlet guide vanes 100 is configured to pre-swirlan airflow 58 provided through the inlet 60 of the nacelle 50, upstreamof the plurality of fan blades 40 of the fan 38. As briefly discussedabove, pre-swirling the airflow 58 provided through the inlet 60 of thenacelle 50 prior to such airflow 58 reaching the plurality of fan blades40 of the fan 38 may reduce separation losses and/or shock losses,allowing the fan 38 to operate with the relatively high fan tip speedsdescribed above with less losses in efficiency.

As discussed, the present disclosure provides a means for reducing icebuildup or ice formation 200 on the inlet pre-swirl feature, e.g.,configured as a plurality of part span inlet guide vanes 100, incommunication with the inlet pre-swirl feature. This provides ananti-icing or de-icing mechanism that prevents the buildup and sheddingof pieces of ice into the engine during adverse weather conditions.

Referring now generally to FIGS. 5A through 8B, in exemplary embodimentsof the present disclosure, the means for reducing ice buildup or iceformation 200 includes a heat source 202 that is in thermalcommunication with the inlet pre-swirl feature, e.g., configured as aplurality of part span inlet guide vanes 100.

Referring to FIG. 5A, a close-up, cross-sectional view of a fan section14 and a forward end of a turbomachine 16 of a turbofan engine 10 inaccordance with an exemplary embodiment of the present disclosure isprovided. The exemplary engine of FIG. 5A may be configured in a similarmanner as the exemplary engine of FIG. 2 described above.

In the exemplary embodiment depicted, the heat source 202 includes anengine bleed airflow 160. For example, in a first exemplary embodiment,referring to FIG. 5A, the heat source 202 includes a bleed air supplyassembly 162 including a bleed air supply duct 164 in airflowcommunication with a high pressure air source 165, e.g., engine bleedairflow 160 from the engine 10, and the leading edge 108 of the inletpre-swirl feature, e.g., configured as a plurality of part span inletguide vanes 100. The bleed air supply duct 164 is configured to receiveand provide the engine bleed airflow 160 to the leading edge 108 of theinlet pre-swirl feature, e.g., configured as a plurality of part spaninlet guide vanes 100. For example, the bleed air supply duct 164 mayprovide the engine bleed airflow 160 to a location proximate the leadingedge 108, e.g., to a location closer to the leading edge 108 than thetrailing edge 110.

In exemplary embodiments, the high pressure air source 165 is thecompressor section, e.g., the LP compressor 22, of the turbomachine 16.For example, in an exemplary embodiment, hot compressor discharge air,e.g., the engine bleed airflow 160, is routed to the leading edge 108 ofthe part span inlet guide vane 100 and then a cooler engine bleedairflow 160 is returned back to the turbomachine 16. It is contemplatedthat the compressor discharge air may be sourced from any stage of theLP compressor 22 or the HP compressor 24.

In this manner, the engine bleed airflow 160 is utilized to heat thepart span inlet guide vane 100 and operate as a means for reducing icebuildup or ice formation at the inlet pre-swirl feature, e.g.,configured as a plurality of part span inlet guide vanes 100.

In an exemplary embodiment, the bleed air supply assembly 162 furtherincludes a bleed air return duct 172 in airflow communication with thebleed air supply duct 164 and a portion of the turbomachine 16. Forexample, the bleed air return duct 172 is configured to return theengine bleed airflow 160 back to the turbomachine 16, e.g., the LPcompressor 22. The return engine bleed airflow 160 may be injected backto the turbomachine 16 at any applicable stage of the turbofan engine 10(e.g., upstream from where the bleed air supply duct 164 received theengine bleed airflow 160) or dumped into the bypass airflow passage 56as described herein.

In some exemplary embodiments, such as the exemplary embodiment of FIG.5A, the bleed air supply assembly 162 further includes a cavity 168 ofthe part span inlet guide vane 100 in communication with a portion ofthe engine bleed airflow 160. Furthermore, the part span inlet guidevane 100 for the embodiment depicted further defines a trailing edgeopening 170, which is in airflow communication with the cavity 168, andthus is in airflow communication with a portion of the bleed air supplyassembly 162. Accordingly, with such a configuration, a portion of theengine bleed airflow 160 may be provided from the bleed air supplyassembly 162 to the cavity 168 of the part span inlet guide vane 100,and further through the trailing edge opening 170 of the part span inletguide vane 100 during operation of the turbofan engine 10 to reduce awake formed by the respective part span inlet guide vane 100.

In an exemplary embodiment, the bleed air supply assembly 162 furtherincludes a valve 166 (depicted in phantom in FIG. 5A) in communicationwith the bleed air supply duct 164. The valve 166 is transitionablebetween an open position in which the bleed air supply duct 164 receivesand provides the engine bleed airflow 160 to the leading edge 108 of theinlet pre-swirl feature, e.g., configured as a plurality of part spaninlet guide vanes 100, and a closed position in which the bleed airsupply duct 164 is not in airflow communication with the leading edge108 of the inlet pre-swirl feature, e.g., configured as a plurality ofpart span inlet guide vanes 100.

Referring now to FIG. 5B, providing a cross-sectional view of a partspan inlet guide vane 100 in accordance with an exemplary embodiment ofthe present disclosure, the locations of the bleed air supply duct 164and the bleed air return duct 172 within the part span inlet guide vane100 are shown. The exemplary part span inlet guide vane 100 of FIG. 5Bmay be configured in a similar manner, as the exemplary part span inletguide vane 100 of FIG. 5A. However, for the embodiment of FIG. 5B, theexemplary part span inlet guide vane 100 does not include a cavity 168or a trailing edge opening 170.

For example, the bleed air supply duct 164 is configured to receive andprovide the engine bleed airflow 160 to the leading edge 108 of the partspan inlet guide vane 100. As shown, the bleed air return duct 172 ispositioned between the bleed air supply duct 164 and the trailing edge110 of the part span inlet guide vane 100. Referring to FIG. 5B, thepart span inlet guide vane 100 includes the leading edge 108, thetrailing edge 110, a pressure side 120, and a suction side 122.

Referring now to FIG. 6 , a close-up, cross-sectional view of a fansection 14 and a forward end of a turbomachine 16 of a turbofan engine10 in accordance with an exemplary embodiment of the present disclosureis provided. The exemplary engine 10 of FIG. 6 may be configured in asimilar manner as the exemplary engine of FIG. 5A described above. Inthe exemplary embodiment of FIG. 6 , the bleed air supply assembly 162further includes a bleed air intermediate duct 174 in airflowcommunication with and disposed between the bleed air supply duct 164and the bleed air return duct 172. The bleed air intermediate duct 174is in airflow communication with an upstream end 176 of the outernacelle 50 and is configured to receive and provide the engine bleedairflow 160 to the upstream end 176 of the outer nacelle 50. In such anembodiment, the engine bleed airflow 160 through the bleed air supplyduct 164 is the hottest portion of the engine bleed airflow 160, theengine bleed airflow 160 through the bleed air intermediate duct 174 isa warm air, and the engine bleed airflow 160 through the bleed airreturn duct 172 is a cooler air return back to the turbomachine 16,e.g., the LP compressor 22. It is also contemplated that in otherexemplary embodiments, the engine bleed airflow 160 could first bedirected to the upstream end 176 of the outer nacelle 50 and then to theleading edge 108 of the part span inlet guide vane 100. It is furthercontemplated that in other exemplary embodiments, the engine bleedairflow 160 could be simultaneously directed to the upstream end 176 ofthe outer nacelle 50 and the leading edge 108 of the part span inletguide vane 100.

In this manner, the engine bleed airflow 160 is utilized to heat thepart span inlet guide vane 100 and the upstream end 176 of the outernacelle 50. As such, the engine bleed airflow 160 is utilized as a meansfor reducing ice buildup or ice formation at the inlet pre-swirlfeature, e.g., configured as a plurality of part span inlet guide vanes100, and at the upstream end 176 of the outer nacelle 50.

Referring now to FIG. 7 , a close-up, cross-sectional view of a fansection 14 and a forward end of a turbomachine 16 of a turbofan engine10 in accordance with an exemplary embodiment of the present disclosureis provided. The exemplary engine 10 of FIG. 7 may be configured in asimilar manner as the exemplary engine of FIG. 5A described above. Inthe exemplary embodiment of the present disclosure depicted, the heatsource 202 includes an engine oil flow or engine oil 210. For example,in another exemplary embodiment, referring to FIG. 7 , the heat source202 includes an oil supply assembly 262 including an oil supply duct 264in flow communication with an oil source 220 of the turbofan engine 10and the leading edge 108 of the inlet pre-swirl feature, e.g.,configured as a plurality of part span inlet guide vanes 100. The oilsupply duct 264 is configured to receive and provide the engine oil 210to the leading edge 108 of the inlet pre-swirl feature, e.g., configuredas a plurality of part span inlet guide vanes 100.

In this manner, the hot engine oil flow 210 is utilized to heat the partspan inlet guide vane 100 and operate as a means for reducing icebuildup or ice formation at the inlet pre-swirl feature, e.g.,configured as a plurality of part span inlet guide vanes 100.

In an exemplary embodiment, the oil supply assembly 262 further includesan oil return duct 272 in flow communication with the oil supply duct264 and the oil source 220. For example, the oil return duct 272 isconfigured to return the engine oil 210 back to the oil source 220.

In an exemplary embodiment, the oil supply assembly 262 may furtherinclude a valve 266 (depicted in phantom) in communication with the oilsupply duct 264. The valve 266 is transitionable between an openposition in which the oil supply duct 264 receives and provides theengine oil flow 210 to the leading edge 108 of the inlet pre-swirlfeature, e.g., configured as a plurality of part span inlet guide vanes100, and a closed position in which the oil supply duct 264 is not inairflow communication with the leading edge 108 of the inlet pre-swirlfeature, e.g., configured as a plurality of part span inlet guide vanes100.

Referring now to FIG. 8A, a close-up, cross-sectional view of a fansection 14 and a forward end of a turbomachine 16 of a turbofan engine10 in accordance with an exemplary embodiment of the present disclosureis provided. The exemplary engine 10 of FIG. 8A may be configured in asimilar manner as the exemplary engine of FIG. 2 described above. In theexemplary embodiment depicted, the heat source 202 includes anelectrical heating element 310 disposed proximate the leading edge 108of the inlet pre-swirl feature, e.g., configured as a plurality of partspan inlet guide vanes 100 (e.g., closer to the leading edge 108 thanthe trailing edge 110). For example, the heat source 202 is morespecifically the electrical heating element 310 disposed in the leadingedge 108 of the inlet pre-swirl feature.

In this manner, the electrical heating element 310 is utilized to heatthe part span inlet guide vane 100 and operate as a means for reducingice buildup or ice formation at the inlet pre-swirl feature, e.g.,configured as a plurality of part span inlet guide vanes 100.

In an exemplary embodiment, the heat source 202 further includes anelectrical supply assembly 362 including an electrical supply cable 364of the turbofan engine 10 in electrical communication with theelectrical heating element 310.

Referring to FIG. 8B, a cross-sectional view of a part span inlet guidevane 100, the locations of an electrical heating element 310 within thepart span inlet guide vane 100 are shown. For example, the electricalheating element 310 includes a first element 312 adjacent the leadingedge 108, e.g., the first element 312 is closer to the leading edge 108than the trailing edge 110, of the inlet pre-swirl feature, e.g.,configured as a plurality of part span inlet guide vanes 100, and asecond element 314 positioned aft of the first element 312 within theinlet pre-swirl feature, e.g., configured as a plurality of part spaninlet guide vanes 100. It is contemplated that one or more electricalheating elements may be used to heat the part span inlet guide vane 100.Referring to FIG. 8B, the electrical heating element 310 may alsoinclude a third element 316, a fourth element 318, and a fifth element320 each positioned aft of the first element 312 within the inletpre-swirl feature, e.g., configured as a plurality of part span inletguide vanes 100.

Referring to FIGS. 9A and 9B, in another exemplary embodiment of thepresent disclosure, the means for reducing ice buildup or ice formation200 includes a vibration assembly 410 in mechanical communication withthe inlet pre-swirl feature, e.g., configured as a plurality of partspan inlet guide vanes 100.

In an exemplary embodiment, the vibration assembly 410 includes apiezoelectric transducer 420 that vibrates the inlet pre-swirl feature,e.g., configured as a plurality of part span inlet guide vanes 100. Itis contemplated that the vibration assembly 410 may include any othervibration devices such as ultrasonic vibration devices. It is alsocontemplated that the vibration assembly 410 may be used with any of theheat source embodiments disclosed in FIGS. 5A-8B.

In an exemplary embodiment, the vibration assembly 410 further includesan electrical supply assembly 462 including an electrical supply cable464 of the turbofan engine 10 in electrical communication with thepiezoelectric transducer 420.

Referring to FIG. 9B, a cross-sectional view of a part span inlet guidevane 100, the location of the piezoelectric transducer 420 within thepart span inlet guide vane 100 is shown. For example, the piezoelectrictransducer 420 is positioned adjacent the leading edge 108 of the partspan inlet guide vane 100.

Furthermore, in another exemplary embodiment, the inlet pre-swirlfeature, e.g., configured as a plurality of part span inlet guide vanes100, is covered with an anti-ice coating 430. For example, in anexemplary embodiment, the anti-ice coating 430 may include an erosioncoating or erosion layer that resists ice accumulation. In an exemplaryembodiment, the erosion coating is polyurethane. In an exemplaryembodiment, the anti-ice coating 430 has a shore hardness between ShoreA50 and Shore D60. In another exemplary embodiment, the anti-ice coating430 has a shore hardness of Shore A90.

It is contemplated that the anti-ice coating 430 may cover the inletpre-swirl feature, e.g., configured as a plurality of part span inletguide vanes 100, as shown in FIG. 9B. However, it is also contemplatedthat the anti-ice coating 430 may only be applied to selected portionsof the inlet pre-swirl feature, e.g., configured as a plurality of partspan inlet guide vanes 100. It is also contemplated that with a partspan inlet guide vane 100 formed of a polymer composite airfoil with ametal leading edge 108, the anti-ice coating 430 is applied only toportions aft of the metal leading edge 108.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

A turbofan engine comprising: a fan comprising a plurality of fanblades; a turbomachine operably coupled to the fan for driving the fan,the turbomachine comprising a compressor section, a combustion section,and a turbine section in serial flow order and together defining a coreair flowpath; a nacelle surrounding and at least partially enclosing thefan; an inlet pre-swirl feature located upstream of the plurality of fanblades, the inlet pre-swirl feature attached to or integrated into thenacelle; and a means for reducing ice buildup or ice formation on theinlet pre-swirl feature, the means in communication with the inletpre-swirl feature.

The turbofan engine of any preceding clause, wherein the means forreducing ice buildup or ice formation comprises a heat source in thermalcommunication with the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the heat source isin communication with a leading edge of the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the heat sourcecomprises an engine bleed airflow.

The turbofan engine of any preceding clause, wherein the heat sourceincludes an air supply assembly comprising a supply duct in airflowcommunication with a high pressure air source and a leading edge of theinlet pre-swirl feature, the supply duct configured to receive andprovide the engine bleed airflow to the leading edge of the inletpre-swirl feature.

The turbofan engine of any preceding clause, wherein the air supplyassembly further comprises a return duct in airflow communication withthe supply duct and a portion of the turbomachine, the return ductconfigured to return the engine bleed airflow back to the turbomachine.

The turbofan engine of any preceding clause, wherein the air supplyassembly further comprises a valve in communication with the supplyduct, the valve transitionable between an open position in which thesupply duct receives and provides the engine bleed airflow to theleading edge of the inlet pre-swirl feature, and a closed position inwhich the supply duct is not in airflow communication with the leadingedge of the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the air supplyassembly further comprises an intermediate duct in airflow communicationwith and disposed between the supply duct and the return duct, theintermediate duct in airflow communication with an upstream end of thenacelle and configured to receive and provide the engine bleed airflowto the upstream end of the nacelle.

The turbofan engine of any preceding clause, wherein the high pressureair source is the compressor section of the turbomachine.

The turbofan engine of any preceding clause, wherein the heat sourcecomprises an electrical heating element disposed in thermalcommunication with a leading edge of the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the electricalheating element comprises a first element adjacent the leading edge ofthe inlet pre-swirl feature and a second element positioned aft of thefirst element within the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the heat sourceincludes an electrical supply assembly comprising an electrical supplycable in electrical communication with the electrical heating element.

The turbofan engine of any preceding clause, wherein the heat sourcecomprises an engine oil.

The turbofan engine of any preceding clause, wherein the heat sourceincludes an oil supply assembly comprising an oil supply duct in flowcommunication with an oil source and a leading edge of the inletpre-swirl feature, the oil supply duct configured to receive and providethe engine oil to the leading edge of the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the oil supplyassembly further comprises an oil return duct in flow communication withthe oil supply duct and the oil source, the oil return duct configuredto return the engine oil back to the oil source.

The turbofan engine of any preceding clause, wherein the oil supplyassembly further comprises a valve in communication with the oil supplyduct, the valve transitionable between an open position in which the oilsupply duct receives and provides the engine oil to the leading edge ofthe inlet pre-swirl feature, and a closed position in which the oilsupply duct is not in flow communication with the leading edge of theinlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the means forreducing ice buildup or ice formation comprises a vibration assembly inmechanical communication with the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the means forreducing ice buildup or ice formation on the inlet pre-swirl feature incommunication with the inlet pre-swirl feature comprises a piezoelectrictransducer that vibrates the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the means forreducing ice buildup or ice formation on the inlet pre-swirl feature incommunication with the inlet pre-swirl feature comprises an anti-icecoating covering the inlet pre-swirl feature.

The turbofan engine of any preceding clause, wherein the inlet pre-swirlfeature comprises a part span inlet guide vane at a location forward ofthe plurality of fan blades of the fan along an axial direction and aftof an inlet of the nacelle.

A turbofan engine comprising: a fan comprising a plurality of fanblades; a turbomachine operably coupled to the fan for driving the fan,the turbomachine comprising a compressor section, a combustion section,and a turbine section in serial flow order and together defining a coreair flowpath; a nacelle surrounding and at least partially enclosing thefan; an inlet pre-swirl feature located upstream of the plurality of fanblades, the inlet pre-swirl feature attached to or integrated into thenacelle; and one or more of: a heat source in thermal communication withthe inlet pre-swirl feature; a vibration assembly in mechanicalcommunication with the inlet pre-swirl feature; a piezoelectrictransducer that vibrates the inlet pre-swirl feature; and an anti-icecoating covering the inlet pre-swirl feature.

This written description uses examples to disclose the disclosure,including the best mode, and also to enable any person skilled in theart to practice the disclosure, including making and using any devicesor systems and performing any incorporated methods. The patentable scopeof the disclosure is defined by the claims, and may include otherexamples that occur to those skilled in the art. Such other examples areintended to be within the scope of the claims if they include structuralelements that do not differ from the literal language of the claims, orif they include equivalent structural elements with insubstantialdifferences from the literal languages of the claims.

While this disclosure has been described as having exemplary designs,the present disclosure can be further modified within the scope of thisdisclosure. This application is therefore intended to cover anyvariations, uses, or adaptations of the disclosure using its generalprinciples. Further, this application is intended to cover suchdepartures from the present disclosure as come within known or customarypractice in the art to which this disclosure pertains and which fallwithin the limits of the appended claims.

1-20. (canceled)
 21. A turbofan engine comprising: a fan comprising aplurality of fan blades; a turbomachine operably coupled to the fan fordriving the fan, the turbomachine comprising a compressor section, acombustion section, and a turbine section in serial flow order andtogether defining a core air flowpath; a nacelle surrounding and atleast partially enclosing the fan; an inlet pre-swirl feature locatedupstream of the plurality of fan blades, the inlet pre-swirl featureattached to or integrated into the nacelle; and an electrical device inthermal communication with the inlet pre-swirl feature, wherein theelectrical device is configured to reduce ice buildup or ice formationon the inlet pre-swirl feature.
 22. The turbofan engine of claim 21,wherein the electrical device comprises an electrical heating element inthermal communication with the inlet pre-swirl feature.
 23. The turbofanengine of claim 22, wherein the electrical heating element is disposedin a leading edge of the inlet pre-swirl feature.
 24. The turbofanengine of claim 23, wherein the electrical heating element comprises afirst element adjacent the leading edge of the inlet pre-swirl featureand a second element positioned aft of the first element within theinlet pre-swirl feature.
 25. The turbofan engine of claim 22, whereinthe electrical device includes an electrical supply assembly comprisingan electrical supply cable in electrical communication with theelectrical heating element.
 26. The turbofan engine of claim 21, whereinthe electrical device comprises two electrical heating elements inthermal communication with the inlet pre-swirl feature.
 27. The turbofanengine of claim 21, wherein the electrical device comprises threeelectrical heating elements in thermal communication with the inletpre-swirl feature.
 28. The turbofan engine of claim 21, wherein theelectrical device comprises four electrical heating elements in thermalcommunication with the inlet pre-swirl feature.
 29. The turbofan engineof claim 21, further comprising a vibration assembly in communicationwith the inlet pre-swirl feature.
 30. The turbofan engine of claim 21,further comprising a piezoelectric transducer that vibrates the inletpre-swirl feature.
 31. The turbofan engine of claim 21, furthercomprising an anti-ice coating covering the inlet pre-swirl feature. 32.A nacelle assembly for a turbofan engine, the turbofan engine comprisinga fan including a plurality of fan blades, the nacelle assemblyconfigured to circumferentially surround the fan, the nacelle assemblycomprising: an inlet pre-swirl feature located upstream of the pluralityof fan blades, the inlet pre-swirl feature attached to or integratedinto the nacelle assembly; and an electrical device in thermalcommunication with the inlet pre-swirl feature, wherein the electricaldevice is configured to reduce ice buildup or ice formation on the inletpre-swirl feature.
 33. The nacelle assembly of claim 32, wherein theelectrical device comprises an electrical heating element in thermalcommunication with the inlet pre-swirl feature.
 34. The nacelle assemblyof claim 33, wherein the electrical heating element is disposed in aleading edge of the inlet pre-swirl feature.
 35. The nacelle assembly ofclaim 34, wherein the electrical heating element comprises a firstelement adjacent the leading edge of the inlet pre-swirl feature and asecond element positioned aft of the first element within the inletpre-swirl feature.
 36. The nacelle assembly of claim 33, wherein theelectrical device includes an electrical supply assembly comprising anelectrical supply cable in electrical communication with the electricalheating element.
 37. The nacelle assembly of claim 32, wherein theelectrical device comprises two electrical heating elements in thermalcommunication with the inlet pre-swirl feature.
 38. The nacelle assemblyof claim 32, further comprising a vibration assembly in communicationwith the inlet pre-swirl feature.
 39. The nacelle assembly of claim 32,further comprising a piezoelectric transducer that vibrates the inletpre-swirl feature.
 40. The nacelle assembly of claim 32, furthercomprising an anti-ice coating covering the inlet pre-swirl feature.